Rocket propulsion is based on the high-speed ejection of propellant mass. Propellant mass, once ejected, does not return and the total mass of the spacecraft plus propellant decreases with each propulsive maneuver. The change in spacecraft velocity ΔV (delta-V) is a function of how much propellant mass Mp was ejected, and the exit speed of that mass with respect to the spacecraft. The rocket equation, given by:ΔV=goIsp ln(Mi/Mf)  (1)relates the change in spacecraft velocity to the specific impulse Isp and the change in total spacecraft mass from an initial Mi to a final Mf. Ejected propellant mass Mp is the difference between Mi and Mf, go is the gravitational acceleration constant at the Earth's surface (9.8 meter/s2), and Isp is the ratio of thrust divided by the mass flow rate. FIG. 1 shows propellant mass fractions Mp/Mi, calculated using Eq. 1, required to reach various velocity increments for several values of specific impulse. The curves are representative of cold gas thrusters (˜50-s), small solid rockets or hydrazine thrusters (˜200-s), bipropellant thrusters (˜300-s), hydrogen/oxygen thrusters (˜450-s), hot hydrogen thrusters (˜900-s), and ion thrusters (˜3000-s).
Cold gas thrusters are the simplest, but are useful for velocity increments below about 300-m/s. Chemical thrusters with specific impulse between 200-s and 450-s are more complex, but they enable significantly lower propellant mass fractions. Chemical thrusters have been the primary workhorses of the Space Age; they regularly launch spacecraft into orbit and have propelled space probes beyond Pluto's orbit. Nuclear and solar thrusters can provide a 900-s Isp with hydrogen propellant, but these have only been demonstrated in ground tests. Electric thrusters top the specific impulse range, but these are typically low-thrust (less than 1−N) devices.
The main reason electric thrusters provide low thrust is that the power required to produce a Newton of thrust increases proportionally with specific impulse. The combination (g0 Isp) is the directed exit speed Ve of the propellant mass; a 200-s Isp thruster, for example, has a directed exit speed of 2.0-km/s while a 3000-s thruster has a directed exit speed of 30-km/s. The kinetic power PKE required to maintain the exhaust plume is proportional to the mass flow rate dm/dt and the square of the directed exit speed Ve:PKE=½dm/dtVe2.  (2)Because thrust T is proportional to mass flow rate times velocity,T=dm/dtVe,  (3)the power per unit thrust is proportional to Ve, and thus, specific impulse Isp.
FIG. 2 shows the power required to generate a Newton of thrust as a function of specific impulse, and the energy density of the propellant in the exhaust stream, assuming complete conversion of input power into directed plume power. Cold gas thrusters utilize propellant thermal energy densities at typical spacecraft temperatures that range from a few tenths to ˜2-MJ/kg. Chemical thrusters use propellants with chemical potential energy densities up to a few tens of MJ/kg. To achieve specific impulses beyond 500-s, the addition of external energy (e.g., thermal or electric power from solar cells or nuclear reactors) to the propellant stream is currently required. Nuclear fuels have potential energy densities that are about a million times higher than chemical propellants, up to tens of TJ/kg. If these could be used directly as propellants, high thrust at 500,000-s Isp or higher would be possible. Presently, however, it is still necessary to rely on low thrust electric propulsion for Isp above 1000-s.
FIG. 2 shows that a megawatt of chemical power is generated by a kilonewton thruster (enough to barely lift the mass of a typical adult male human at the Earth's surface) at 200-s Isp. The U.S. Space Shuttle solid rocket motors generate 30 gigawatts of power to generate a total thrust of 25-MN at 242-s Isp, which is equivalent to the average instantaneous electrical power usage for the entire state of California.
A significant propulsion challenge is to land humans or equipment on the moon and return them to Earth. The velocity increment for this mission about 11,500 m/sec. This is a very large velocity increment for an in-space mission.
Sending a spacecraft to geosynchronous Earth orbit (GEO) from an initial 400 km circular low Earth orbit (LEO) has a minimum propellant mass fraction of 83% using 220 sec Isp thrusters and 59% using 440 sec Isp thrusters.
For the 400 km LEO to moon landing and return scenario with a ΔV of 11,500 msec, the propellant mass fraction is 93% using the best chemical thruster Isp of 440 sec. The only way to accomplish this mission using chemical thrusters is to use multiple stages and throw away unneeded mass such as empty propellant tanks whenever possible. The Apollo moon landings of the late 1960s and early 1970s, for example, started with 118,000 kg in a LEO parking orbit and put 47,000 kg in low lunar orbit using a 400 sec Isp thruster. The remaining maneuvers used space-storable propellants with a reduced Isp of 315 sec. The mission ended with a 5800 kg capsule entering Earth's atmosphere at near escape velocity. Less than 5% of the original mass in LEO was returned.
Dramatic reductions in required propellant mass occur as Isp is increased. Unfortunately, electric thrusters are required to attain higher Isp, and the input energy requirements, for constant thrust, scale with Isp. This is why electric thrusters with high Isp generally have low thrust levels. A 1000 N (224 lb-force) thruster, for example, would require 30 MW of electric power at 5000 sec Isp, and 60 MW at 10,000 sec Isp. As Isp increases, the propellant mass savings is eliminated at some point by an increasing mass of the power supply.
It would be useful to be able to provide a spacecraft propulsion technology in which the amount of propellant mass required (e.g., to facilitate a particular spacecraft maneuver) is reduced. It would also be useful to be able to provide a spacecraft propulsion technology in which propellant mass and/or energy is used more efficiently.
It would be useful to be able to provide propulsion systems and methods that are more mass and/or energy efficient than conventional spacecraft propulsion technologies.
It would be useful to be able to continually use or reuse, rather than discard, expensive spacecraft components.